Fatigue Life Analysis of Al 8090 Helicopter Fuselage Panels

Abstract:

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Considering the aerospace structures, the advantages of Al-Li alloys in comparison with
conventional aluminium alloys comprise relatively low densities, high elastic modulus, excellent
fatigue and toughness properties, and superior fatigue crack growth resistance. Unfortunately, these
alloys have some disadvantages due to highly anisotropic mechanical properties and due to a very
high crack growth rate for microstructurally short cracks. This could mean relatively early cracking
in high stress regions such as rivet holes in helicopter fuselage panels. Consequently a more
accurate approach in fatigue life analysis is requested. Considering that the 8090 T81 aluminium
alloy has been widely used in an helicopter structure, in particular in the bolted connection between
the stringers and the modular joint frame in the rear of the fuselage, it is extremely important to
found a reliable procedure for the fatigue life assessment of the component. Thus, using the results
of experimental tests made on panel specimens, a FE general model and two submodels of the
critical zone (involved in fatigue damage during the tests) have been modelled in order to
investigate the complex state of stress near the rivets holes. These stress values obtained have been
elaborated for a fatigue assessment.

Abstract: A two-parameter constraint-based fracture mechanics approach is used to explain the effect of the constraint on the apparently anomalous behavior of short fatigue cracks. The different levels of stress constraint are quantified by the T-stress, and microstructurally as well as mechanically short cracks are discussed. Short cracks generally behave more sensitively to the constraint than the long ones. It is shown that in most cases, the existence of short cracks goes hand in hand with an intrinsic loss of the constraint, which contributes to a decrease of their fatigue
threshold values and accelerates their growth. In this paper, the above effect is quantified and conclusions concerning the applicability of the fracture mechanics parameters and approaches to the estimation of the residual fatigue life of structures are discussed.

Abstract: In this paper, the mechanisms of propagation of the damage in aluminum panels repaired with bonded composite patches of different mechanical characteristics is analyzed. The aim of this study is to analyze analytically, experimentally and numerically the advantage of the use of bonded composite patches to increase the fatigue life and to reduce the state of tension at the crack tips. The experimental results show that both static strength and fatigue life of the repaired aluminum panel has significantly increased due to the bonded composite patches. The different patches and adhesive, used for cracked panels, have provided about a 100-110% improvement in the fatigue life and a 30-35% decrease in the stress intensity factor. A comparison between finite elements calculations and experimental data has been carried out. The good agreement between the experimental data and the numerical ones has demonstrated the possibility to obtain an optimized design of bonded patches with the numerical tools.

Abstract: In this research, fatigue behaviour of Ti-6Al-4V alloy was investigated for smooth and notched specimens with stress concentration factor(Kt) 3.6 and 4.1.This investigation was conducted for various diameter bars having different ultimate strength.Rotating bending fatigue test at R= -1 was emploied for this research. Notch sensitivity data was compared with those of steels. The result indicated that the presence of notch in this alloy has a different amount of sensitivity when the notch specimens were subjected to high cycle fatigue (HCF) and low cycle fatigue(LCF) tests.The notch sensitivity of this alloy was shown generally to be much lower than steel alloys with similar ultimate strength values. Therefore,considering the low sensitivity to notch of this alloy, can be recommended for applications with the presence of notch such as biomedical application

Abstract: Al-Li alloy 2198-T8 was used in the fuselage application. Integral fuselage panels were joined by double friction stir welds. Fatigue tests were conducted in the R=0.1. Notch was made between two welds. Residual stresses were measured and analyzed in the test samples with double welds. Cracks grew from the centre of two welds and across the two welds were observed, and crack growth rates were measured and compared with parent material. It is shown that crack growth rates are lower between double welds, and it is close to parent material after cross the two welds. The virtual crack closure technique (VCCT) method was used to calculate stress intensity factor from residual stress (Kres) in aim to explain the experimental findings.

Abstract: Currently, self-piercing riveting (SPR) is a major technology used by manufacturers to join aluminium body structures to reduce vehicle weight. Normally, for SPR of one specific stack more than one die, rivet, and velocity combination can be applied. Which parameter combination is chosen is depending on the surrounding joints. In order to increase productivity and reduce the number of robots used, it is preferred to use the same rivet/die combination for as many joints as possible. This means for the same stack, different die may be used. To see the influence of die profiles on joint quality, a DF die, which would generate severe cracks and a DC die, which would generate no cracks or only small cracks, were used to join two stacks with a high strength aluminium alloy, AA6008T61, as the bottom layer. The joint quality was analyzed, and the static and fatigue strengths of these stacks were studied. Results showed that cracks on joint buttons might reduce static and fatigue lap shear strength but had no obvious influence on static and fatigue T peel strength for the joints studied.